Homing system for guided missiles



Nov. 23, 1965 A. wr-:LTl

HOMING SYSTEM FOR GUIDED MISSILES 2 Sheets-Sheet 1 Filed Dec. 6, 1961Nov. 23, 1965 Filed Dec. 6, 1961 A. wELTl 3,219,294

HOMING SYSTEM FOR GUIDED MISSILES 2 Sheets-Sheet. 2

Eslevaton Z eI'VO l Elevatlon l Resolver tl- A' RESOLVER 0 SYSTEM ANALOGHEIGHT l 0c, COMPUTER CONTROL l l MOTOR L Detector l FTA-)T g1 SERVOS iLateral l Sub- AA 1 dt Lateral Servo 3 Resolver nach? dt E H ServoControl NET NE-rwo RK) 24 cgrgllrRLOI-l WQRK l "/MOTOR servo AmiCOMPUTER 23- L@ Lateral cumrol STAGE N- l Servo ,4 il Lateral l D tResover ,g` A A $2 e ec or Sub AR i d t l I C tract dt l f/3 2 NET 1 NET J TS|DE im L5 TI z 4 sERvOs L L6 RESOLVER l Elevation SYSTEM Elevahon 'rResolver Servo Mfg f1 IMAGE AREA scANNER "GYRO 4 FIELD LENS a r 5 6?-RAD|AT|0N RESPONSIVE CELL OBJECT LENS `4 g *To SERVO 8 Fig` 5 TO SERVOUnited States Patent O 3,219,294 HOMHNG SYSTEM FUR GUIDED MISSILES ArnoWelti, Zurich, Switzerland, assigner to Albiswerk Zurich A.G., Zurich,Switzerland, a Swiss corporation Filed Dec. 6, 1961, Ser. No. 157,506Claims priority, application Switzerland, Dec. 7, 1960, 13,665/60 4Claims. (Cl. 244-14) My invention relates to a method and system forhoming guidance of missiles, according to which a missile, hereinunderstood to denote any object travelling through space and providedwith suitable directional control means such as control surfaces orcontrol jets, guides itself toward a target, which may likewise betravelling in space, to bring about a collision or near-collision ofmissile and target. In a more particular aspect, my invention relates toa homing method and system of the passive type in which energy from thetarget itself, for example heat radiation, is sensed by suitabledetectors, for example infrared sensors, in the missile, for therebytracking the target and determining and `correcting any departures fromthe collision course. The control signals for actuating the controls,for example a rudder, to produce the required course correction, aredetermined by means of computers.

Such determination of the control signals in the known homing systems ispredicated upon the collision criterion that the geometric angle betweenthe line of sight C from the missile F to the target Z on the one hand,and a spacially fixed reference axis K on the other hand must remaininvariable in time. Referring to a missile-fixed polar coordinate systemaccording to FIG. 1 of the accompanying drawings, this angle deiines theso-called spherical distance d whose relation to the angular coordinatesof the line of sight C and the stable axis K- namely the lateral vangleal and the elevation k1 of C, and the lateral angle a2 and elevation A2of K- is expressed by the equation:

cos d=sin )q sin z-i-cos k1 cos k2 cos (u2-a1) The equipment forcomputing collision-course errors on the basis of this equation isintricate and spacedemanding, this being a considerable disadvantage ofthe known homing-guidance systems. Furthermore, exacting accuracyrequirements are to be met by the angle measurements, which likewiserequires complicated and expensive apparatus.

It is an object of my invention to minimize or obviate thesedisadvantages.

My invention is based upon the following considerations:

A missile moving in space and a likewise moving target will collideunder all circumstances if the target image remains at standstill on theviewing or picture area of the detector in the missile, it being onlypresumed that the collision has not yet occurred and that target andmissile do not travel parallel to each other. The immobility of thetarget image is tantamount to the fact that initially the line of sightfrom missile `to target retains its direction in space constant duringthe interval of time under observation. Consequently, if the missile iscontrolled so that the direction-al invariance of the line of sight ispreserved at any moment, the impact condition is also satisfied at anymoment.

Such directional invariance of the line of sight can be controlled aftermeasuring its direction with the aid of target tracking, and referringit to the direction of a spacially fixed axis relative to a polarcoordinate system fixed with respect to the missile. The impactcondition is met, according to the basic concept of the invention,

3,2l994 Patented Nov. 23, 1965 if the dilference of correspondingangular coordinates of these two directions is kept constant in time. Itshould be understood that this is not identical with determining theabove-mentioned spherical distance d between the line of sight and thestable reference axis; the above-mentioned directions are rather definedby their angular coordinates in the sense of lateral and elevationalangles in the polar coordinate system of the missile, and it is onlysignificant that the difference of the lateral angles and the differenceof the height angles remain constant.

The impact condition, thus formulated, can be graphically represented inthe so-called phase space, a diagram which represents the condition oftwo magnitudes in mutual correlation. According to FIG. 2 of thedrawing, the angle differences Aa and Ak at an image point P arecorrelated as the two Cartesian coordinates of that point. The impactcondition is satisfied if the image point P remains at standstill in thephase space.

Disturbances due to extraneous influences may cause the missile todepart from its collision course. In the phase such departure manifestsitself in that the image point P commences to wander. However, as soonas the image point, under the action of the control perfor-mance in themissile, again comes to standstill, the impact condition is againsatisfied. Although during the transition interval the direction of theline of sight in space has changed, this merely delays, but does notprevent, the collision. Consequently, the impact condition does notconstitute a rigid geometrical requirement but has a tempora-rycharacter. Each departure from the collision course, due to disturbance,causes a control action in the missile with the effect of slowing theimage-point motion in the phase space down to standstill, and theninitiates a new collision course of the missile. In other words, thecollision course is reset upon each such disturbance. This ight-controlprinciple diffe-rs fundamentally from that involved in the known systemsin which the control system takes care that the collision courseinitially followed by the missile is retained up to impact with thetarget.

The impact condition upon which the control method of the invention isbased, namely does not quite fully satisfy the above-mentioned conditionaccording to which the spherical distance d between lthe line of sightand the .stable axis is to remain const-ant. Applicable is the equation:

Accordingly, the spherical distance d would be constant only as long as,aside from the angle difference Aa and AA, the angle A2 also remainedconstant.

Although the prerequisites A=constant and A t=con stant are suicient asimpact condition, a limitation in the freedom of motion of the missileis necessary. One reason for such limitation is the fact that thedetermination of the direction C and K (line of sight, and spacial- 1yfixed axis respectively) are also subject to limitation, for example,due to the constructionally limited turning range of a radiation sensinghead, or the frame or gimbal stops of a gyro system. Furthermore, themissile is supposed to take the course of most favorable impactprobability. This requires roll-motion stabilization of the missile. Inaddition, the missile is to be so controlled as to minimize the durationof the transition interval during which the missile passes from one toanother collision course upon occurrence of course disturbance. Thismeans, relative to the phase space, that the return of the image pointto a standstill position again satisfying the impact condition shouldalways occur on a straight path.

A stability condition satisfying those requirements can be derived fromthe phase space shown in FIG. 3. Assume that the image point hasmigrated from the initial standstill point P to a point P1 due tolateral deflection of the missile. The directional regulator is nowcalled upon to actuate the control surfaces or other control means insuch a manner that the error from P0 to P1, after termination of atransitional interval, is corrected by placing the target image onto anew standstill point P2. It must be taken into consideration that thecomponents Aa and AA of the direction regulator for a skewed position ofthe instantaneous missile course to the target course, are dependentupon each other, this being the reason why there is the danger of theso-called overlapping of coordinates in the regulating and controlsystem which may lead to undesired hunting motion with respect toelevational and lateral positions of the missile. Under such conditions,a satisfactory control of the missile would be infeasible. Y'Instabilities of this kind, as are inevitably expectable with a motionof the image point in the phase space on a curved path such as the oneindicated by a broken line in FIG. 3, can be avoided by a return on astraight path. The condition for such straight return is that thedirectional angle in the phase space must remain constant. Consequently,the following condition applies:

fKAoz) tan constant 061-062 and l-*Z and of the ratio both down to thezero value. As explained above, the terms al and M denote the angularcoordinates of the sighting line C from the missile to the target, anda2 and A2 denote the angular coordinates of the stable axis K, theseangular coordinates being continuously measured relative to a polarcoordinate system fixed with respect to the missile.

The apparatus according to the invention for performing themissile-guiding method explained above, comprises a gyro to provide astable reference axis and a radiation detector for tracking the target.These two components control respective angle transmitters or resolverswhich furnish the instantaneous llateral and elevational angles of theline of sight and of the stable axis respectively with respect to apolar coordinate system fixedly related to the missile. The apparatusfurther comprises two networks for producing two control signals whichform a measure of the time derivation (rate of change) of the differencebetween the lateral angles on the one hand, and of the time derivation(rate of change) of the difference between the elevational angles on theother hand; and these two signals control respective motors correlatedto two coordinate axes of the flight directional control means in thesense of reducing the rate of angular change. The system furthercomprises a differential amplifier to which one of the above-mentionedtwo control signals is directly supplied and to which the other controlsignal is supplied through a regulating amplifier whose gain isregulated in dependence upon the output signal of the differentialamplifier in the sense toward reducing this output signal, and theadjusting member for gain regulating simultaneously controls an angletransmitter (resolver system) whose output signal constitutes a measureof the instantaneous amplification factor of the regulating arnplifierand, upon differentiation, controls a motor correlated to a third axisof the iiight directional control means toward reducing the change ofthis output signal.

According to another feature of the invention, the above-describedhoming system is provided with another radiation detector whichcooperates with the gyro for tracking a given point of the stablereference axis to thereby control the one appertaining resolver.

The invention will be further explained with reference to an embodimentof missile guidance control apparatus illustrated by way of example inthe accompanying drawings in which:

FIGS. l, 2 and 3 are explanatory diagrams already described above;

FIG. 4 is a block diagram of a missile-borne homing control apparatus,according to the invention;

FIG. 5 shows schematically the basic design -of an infrared radiationdetector employed in the embodiment according to FIG. 4; and

FIG. 6 is a circuit diagram of electrical equipment forming part of thesame apparatus.

The guidance apparatus according to FIG. 4 is provided with a radiationdetector 1 for tracking the target Z. The detector 1 is horizontally andvertically rotatable. A set of servomotors 2 is connected with thedetector unit for directing the sensing axis onto the target Z. Thedetector 1 furnishes an error voltage which constitutes a measure forthe departure of the optical axis from the line of sight C. This errorvoltage is impressed upon a servo-control stage 3 where it is convertedto control signals for the servos 2 in the sense required for reducingthe above-mentioned departure. Servo devices of this type for automatictracking of a target are known as such. They are described for examplein the 1959 Proceedings of the IRE, pages 1577-1581, in an articleentitled Servomechanisrns Design Considerations for Infrared TrackingSystems, and in the June 1960 issue of Space/Aeronautics, pages 169etc., in an article entitled Firestreak is Guided by Advanced IR Homer.Also such systems are described in United States Patents No. 2,931,912and No. 2,961,545. However, reference to FIG. 5 will be had for furtherexplaining the operation of the radiation-responsive detector.

According to FIG. 5, the radiation detector comprises an object lens 4,an image-area scanner 5, a field lens 6, and a radiation-responsive cell7. The optical and scanning components are accommodated within a housing4a mounted on a supporting structure which comprises two shafts 8 and 9extending perpendicular to each other and permitting the optical axis ofthe detector to be turned about the respective shaft axes. The shafts 8and 9 are connected with respective servomotors in the set 2 (FIG. 4).The scanner 5 is a rotating raster disc with a raster pattern sodesigned that the beam of light impinging upon the photocell 7 ismodulated in dependence upon the locality of the target-image point onthe image area. Hence the output voltage of the cell 7 is modulated inthe same manner. This cell voltage is compared with a reference voltageindicative of the direction of the optical axis of the radiationdetector and derived from the turning motion of the radiation detector.As long as the image point of the target coincides with the opticalaxis, the result of the comparison is zero and no error voltage isproduced. When the image point of the target moves away from the opticalaxis, the voltage comparison results in a finite error voltage which iseffective in the servo-control unit 3 to cause actuation of theservomotors in the sense of eliminating the departure of the targetimage point from the optical axis. More cornplete details can be hadfrom the above-mentioned references.

Mechanically coupled with the radiation detector 1 is an electric angletransmitter (resolver system) 10 (FIG.

4) which has two output channels. One channel furnishes a voltageproportional to the lateral angle a1. The other output channel furnishesa voltage proportional to the elevation angle A1. The angles a1, A1represent the coordinates of the line of sight C relative to amissilefixed polar coordinate system.

The lateral angle al of the missile target sighting line, the lateralangle a2 of the stable reference axis relative to a missile-fixed polarcoordinate system, the elevation angle A1 of the missile target sightingline and the elevation angle A2 of the stable reference axis relative tothe missile-fixed polar coordinate system may be continuously measuredin any suitable manner known. There are a number of known methods forcontinuously measuring the angular displacement of an axis.

The space-fixed axis K is constituted by the main axis of a gyro system11. The position of the gyro axis relative to the missile-fixed polarcoordinate system can be determined by ascertaining a point of the gyroaxis. In the illustrated embodiment, a light source 12 is used for thispurpose. The source is mounted at the end of the gyro shaft and actsupon a radiation detector 13. The detector 13 may be of the same type asthe detector 1 used for tracking the target, and is likewise equippedfor automatic tracking of the light source 12. For this purpose, thedetector 13 is connected with a servo-control stage 14 and a servomotorset 15 which correspond to the respective devices 3 and 2. Anydirectional change of the missile axis manifests itself in a departureof the image point produced by the light source 12, from the opticalaxis of the radiation detector 13. The resulting corrective motion istransmitted to the resolver system 16 mechanically coupled with theradiation detector 13. The resolver system 16, operating in the samemanner as the resolver system 10, furnishes two output voltagesproportional to the angular coordinates a2, A2 of the stable referenceaxis inthe missile-fixed polar coordinate system.

The described method and means for determining the position of a freegyro axis have the advantage that it requires no force ortorque-demanding angle transmitters at the gyro itself. In lieu of alight source, any other electromagnetically detectable reference point(active or passive radiation source or sink) may be provided on the gyroshaft and a correspondingly sensitive cell in the radiation detector.

The portion enclosed in FIG. 4 by a dot-and-dash line and denoted byA.C. constitutes a regulating system interposed between theabove-described two groups of selftracking sensing components and theflight-direction control motors properof the missile. In the presentembodiment, the intermediate system A.C. is essentially an analogcomputer which determines from the measured input magnitudes a1, a2, A1and A2, the control magnitudes required for flight control. The computerportion A.C. is equipped with two networks 17 and 18 for forming theangle difference AAzAl-Ag and trazar-a2 respectively. The computerfurther comprises two networks 19 and 20 for forming the differentialquotients d(AA)/dt and KAM/dt. The control magnitudes supplied from thenetworks 19 and 20 are supplied to respective control motors 21 and 22for actuating the elevational control (H) and lateral control (L) of themissile. These two control magnitudes also pass into a computer stage 23in which a control signal is generated proportional to the time changein the ratio of the mentioned two control magnitudes, and the lattercontrol signal is applied toy a control motor 24 for actuating the rollcontrol (Q) of the missile.

FIG. 6 illustrates the circuit diagram of the regulating computerportion A.C. according to FIG. 4. The volt-ages UA1, UA2 Ual and Uazsupplied from the resolvers and 16 (FIG. 4) and impressed uponrespective pairs of input terminals 25, 26, 27 and 28, are indicative ofthe continuously measured angular coordinates A1, A2, al

and a2 respectively. The voltages, for example, may be linearlyproportional to these coordinates. Connected across the pairs ofterminals are respective resistors 29, 30, 34 and 35. The voltages UA1and UA2, impressed across respective resistors 29 and 30 areseries-opposed to each other so that the ydifference voltage UAA:UA1-UA2 is obtained across the series connection of the two resistors 29and 3f). Connected to the series connection is a differentiating memberconsisting of a longitudinal capacitor 31 and a transverse resistor 32.Across the resistor 32 there appears a voltage UH proportional to thedifferential quotient d(AA)/dt. The voltage UH is -impressed upon outputterminals 33 where it is available as control magnitude for elevationalcontrol. That is, this voltage UH serves to control the operation ofmotor 21 (FIG. 4) in the event a change in elevational angle isnecessary for returning the missile to a collision course.

Analogously, the resistors 34 and 35 connected across respectiveterminal pairs 27 and 28 are impressed by the voltages Ual and Ua2 inmutally series-opposed relation so that the series connection of the tworesistors 34 and 35 furnishes the difference voltage UAazUal-Uaz. Adifferentiating member is connected to the series connection andconsists of a longitudinal capacitance member 36 and a transverseresistance member 37. The voltage Us appearing across the resistancemember 37 is proportional to the differential quotient d(Az)/dt. ThisVoltage is impressed across terminals 38 for lateral control of themotor 22 (FIG. 4).

The regulating portion A.C. of the system further comprises adifferential amplifier 39 which receives the control voltage UHdirectly, and which is supplied with the control voltage US through avariable regulating amplifier 40. Thus, the output of the regulatingamplifier 40, and one input to the amplifier 39, is equal to the inputvoltage Us to the amplifier 40 times its amplification V; `or theproduct USV. Connected to the output leads of the differential amplifier39 is an actuator or motor 41 having a shaft 42. The angle of rotationof the shaft 42 is designated The shaft 42 is coupled with a controlmember for varying the gain in the regulating amplifier 40 so that theamplification V of the regulating amplifier 40 is inversely proportionalto the rotational angle This accomplishes an inverse feedback from theamplifier 39 through the motor M and the amplifier 40 so as to reducethe output voltage at the differential amplifier 39. This equalizes theinput voltages to the amplifier. Thus But as stated, the amplification Vis controlled to be inversely proportional to so that l VN- Substitutingfor V in the equation USV: UH we obtain l Us?- Un and Us NH Substitutingfor Us and UH in the equation US/ UH We obtain .Connecting adirect-voltage source 43 furnishing a constant voltage E, is apotentiometer rheostat 44 whose slider 45 is mounted on the shaft 4Z sothat the slider 45 taps olf the voltage E proportional to the rotationangle of the shaft 42. Connected to the slider and the potentiometer isa differentiating member, consisting of a longitudinal capacitor 46 anda transverse resistor 47. This differentiating member forms the timederivation (rate of change) of the voltage E, whereby the voltage acrossthe output terminals 48 is proportional to the differential quotient:

d(Aa)/dt (MAM/dt dt This output voltage is applied to control motor 24(FIG. 4) and serves as a control magnitude for the roll control of themissile.

The homing method and ight regulating system according to the inventionaffords a number of advantages. The means for performing the method,involving a zero principle, are relatively simple and require relativelyfew and rather compact components, in comparison with known homingmethods and systems. Furthermore, at suflicient missile velocity, thereliability of impact with the target is considerably increased. This isbecause the performance of the commands issuing to the directionalcontrol means of the missile is not affected =by impairment andinaccuracy or aging of the optical, mechanical or electrical components,including the sensors, as well as by any eccentrical motion of thedriving means, or the occurrence of angular departures between thelongitudinal missile axis and the tangent of the travel course. Thetrajectory or collision course of the missile is not predetermined. Oneach disturbance, irrespective of its character, the flight guidancesystem sets the missile to a new impact course, thus always providing anew definition of the path of light. The high degree of precisionrequired for remote control operation, is not necessary, and thetolerance limits can be kept conveniently wide. All regulating,controlling and computing operations are performed with relativemagnitudes only. As a result, the invention affords an adaptation of themissile flight control to the intended purpose to an extent far beyondthat heretofore attainable.

I claim:

1. A missile-borne system for homing a guided missile onto a target,comprising directional control means having a lateral control motor forcontrol about a lateral axis of the missile, a vertical control motorfor control about a vertical axis of the missile and a longitudinalcontrol motor for control about a longitudinal axis of the missile;gyroscopic reference -means having a stable axis relative to space andindicating means for indicating said stable axis; a radiation sensorresponsive to radiation from the target directed to receive radiationfrom said target for detecting departure of the sensor axis from themissile-target sighting line; detecting means directed at the indicatingmeans of said gyroscopic reference means for detecting departure of saidstable axis from a missile-fixed reference position; two servomotormeans of which one is connected with said sensor for causing it to trackthe target and the other is connected to said detecting means forcausing it to track said stable axis; two angle-transmitting revolvingmeans connected to said respective radiation sensor and detecting meansto issue respective pairs of coordinate voltages of which one voltagecorresponds to a lateral angle and the other to an elevation angle in amissile-fixed polar coordinate system, one pair of coordinate voltagesrelating to the missile-target sighting line and the other pair to saidstable axis; and a computing system comprising two electric networksconnected to said two resolvers and having respective output signalvoltages corresponding to the rate of change of the difference betweensaid two lateral angles and to the rate of change of the difference'between said two elevation angles, said two signal voltages beingconnected to said lateral control motor and to said vertical controlmotor for lateral and vertical control respectively in the senserequired for reducing said rates of change respectively, a differentialamplifier having two input circuits and an output circuit, one of saidinput circuits being connected to one of said signal voltages, avariable-gain regulating amplifier connecting said other signal voltagewith said other input circuit, gain regulating means connected with saidoutput circuit and in controlling connection with said regulatingamplifier for controlling its gain in dependence upon the amplifiedoutput voltage of said differential amplifier and in the sense towardreducing said latter voltage, an angle transmitter connected with saidgain regulating means to be controlled thereby and having an outputsignal indicative of the instantaneous amplification gain of saidregulating amplifier, and a differential-forming network connectedbetween said angle transmitter and said longitudinal control motor forcontrolling the said longitudinal control motor in the sense required toreduce said output signal.

2. A missile-homing system, as claimed in claim 1, wherein said gainregulating means comprises an actuator electrically connected with saiddifferential amplifier output circuit and mechanically connected withsaid regulating amplifier for varying its gain, said angle transmittercomprising a source of constant direct voltage, a potentiometerconnected across said source and having a slide contact, said slidecontact being mechanically connected with said actuator to be displacedthereby in accordance with changes in gain of said regulating amplifier,said differential-forming network being electrically connected to saidslide contact and said source.

3. A missile-homing system, as claimed in claim 1, wherein said stableaxis is constituted by a shaft of said gyroscopic reference means, saidmeans for detecting departure of said stable axis from a missile-fixedreference position comprising a radiation member on said shaft, and asensor responsive to said member and connected to said other servomotormeans for tracking said member.

4. A missile-borne system for homing a guided missile onto a target,comprising directional control means having a lateral control motor forcontrol about a lateral axis of the missile, a vertical control motorfor control about a vertical axis of the missile and a longitudinalcontrol motor for control about a longitudinal axis of the missile;gyroscopic reference Imeans having a stable axis relative to space andindicating means for indicating said stable axis; a radiation sensorresponsive to radiation from the target directed to receive radiationfrom said target for detecting departure of the sensor axis from themissile-target sighting line; detecting means directed at the indicatingmeans of said gyroscopic reference means for detecting departure of saidstable axis from a missile-fixed reference position; two servomotormeans of which one is connected with said sensor for causing it to trackthe target and the other is connected to said detecting means forcausing it to track said stable axis; two angle-transmitting revolvingmeans connected to said respective radiation sensor and detecting meansto issue respective pairs of coordinate voltages of which one voltagecorresponds to a lateral angle al and a2 respectively and the other toan elevation angle 1 and X2 respectively in a missile-fixed polarcoordinate system, one pair of coordinate (al, k1) voltages relating tothe missile-target sighting line and the other pair (a2, x2) to saidstable axis; and a computing system comprising two electric networksconnected to said two resolvers and having respective output signalvoltages corresponding to the rate of change of the difference Ao=a1a2Ibetween said two lateral angles and to the rate of change of thedifference A= \1-)\2 between said two elevation angles, said two signalvoltages being connected to said lateral control motor and to saidvertical control motor for lateral and vertical control respectively inthe sense required for reducing said rates of change d(Au)/dt andd(A)\)/dt respectively, a differential amplifier having two inputcircuits and an output circuit, one of said input circuits beingconnected to one of said signal voltages, a variable-gain regulatingamplifier connecting said other signal voltage with said other inputcircuit, gain regulating means connected with said output circuit andsaid regulating amplifier, and a differential-forming network connectedbetween said angle transmitter and said longitudinal control motor forcontrolling the said longitudinal control motor in the sense required toreduce in controlling connection with said regulating amplifier 5 SaidOutput signal' for controlling its gain in dependence upon the amplifiedoutput Voltage of said differential amplifier and in the sense towardreducing said latter voltage, an angle transmitter connected with saidgain regulating means to be controlled thereby and having an outputsignal proportional to indicative of the instantaneous amplificationgain of References Cited bythe Examiner UNITED STATES PATENTS 2,557,4016/1951 Agins et al. 244-14 10 2,992,423 7/1961 Floyd et al. 244-14 X3,005,981 10/1961 Sanders et al. 244--14 X BENJAMIN A. BORCHELT, PrimaryExaminer.

5 SAMUEL FEINBERG, SAMUEL ENGLE, CHESTER L. JUSTUS, Examiners.

1. A MISSILE-BORNE SYSTEM FOR HOMING A GUIDED MISSILE ONTO A TARGET,COMPRISING DIRECTIONAL CONTROL MEANS HAVING A LATERAL CONTROL MOTOR FORCONTROL ABOUT A LATERAL AXIS OF THE MISSILE, A VERTICAL CONTROL MOTORFOR CONTROL ABOUT A VERTICAL AXIS OF THE MISSILE AND A LONGITUDINALCONTROL MOTOR FOR CONTROL ABOUT A LONGITUDIAL AXIS OF THE MISSILE;GYROSCOPIE REFERENCE MEWANS HAVING A STABLE AXIS RELATIVE TO SPACE ANDINDICATING MEANS FOR INDICATING SAID STABLE AXIS; A RADIATION SENSORRESPONSIVE TO RADIATION FROM THE TARGET DIRECTED TO RECEIVE RADIATIONFROM SAID TARGET FOR DETECTING DEPARTURE OF THE SENSOR AXIS FROM THEMISSILE-TARGET SIGHTING LINE; DETECTING MEANS DIRECTED AT THE INDICATINGMEANS OF SAID GYROSCOPIC REFERENCE MEANS FOR DETECTING DEPARTURE OF SAIDSTABLE AXIS FROM A MISSILE-FIXED REFERENCE POSITION; TWO SERVOMOTORMEANS OF WHICH ONE IS CONNECTED WITH SAID SENSOR FOR CAUSING IT TO TRACKTARGET AND THE OTHER IS CONNECTED TO SAID DETECTING MEANS FOR CAUSING ITTO TRACK SAID STABLE AXIS; TWO ANGLE-TRANSMITTING REVOLVING MEANSCONNECTED TO SAID RESPECTIVE RADIATION SENSOR AND DETECTING MEANS TOISSUE RESPONSIVE PAIRS OF COORDINATE VOLTAGES OF WHICH ONE VOLTAGECORRESPONDS TO A LATERAL ANGLE AND THE OTHER TO AN ELEVATION ANGLE IN AMISSILE-FIXED POLAR COORDINATE SYSTEM, ONE PAIR OF COORDINATE VOLTAGESRELATING TO THE MISSILE-TARGET SIGHTING LINE AND THE OTHER PAIR TO SAIDSTABLE AXIS; AND A COMPUTING SYSTEM COMPRISING TWO ELECTRIC NETWORKCONNECTED TO SAID TWO RESOLVERS AND HAVING RESPECTIVE OUTPUT SIGNALVOLTAGES CORRESPONDING TO THE RATE OF CHANGE OF THE DIFFERENCE BETWEENSAID TWO LATERAL ANGLES AND TO THE RATE OF CHANGE OF THE DIFFERENCEBETWEEN SAID TWO ELEVATION ANGLES, SAID TWO SIGNAL VOLTAGES BEINGCONNECTED TO SAID LATERAL CONTROL MOTOR AND TO SAID VERTICAL CONTROLMOTOR FOR LATERAL AND VERTICAL CONTROL RESPECTIVELY IN THE SENSEREQUIRED FOR REDUCING SAID RATES OF CHANGE RESPECTIVELY, A DIFFERENTIALAMPLIFIER HAVING TWO INPUT CIRCUITS AND AN OUTPUT CIRCUIT, ONE OF SAIDINPUT CIRCUITS BEING CONNECTED TO ONE OF SAID SIGNAL VOLTAGES, AVARIABLE-GRAIN REGULATING AMPLIFIER CONNECTING SAID OTHER SIGNAL VOLTAGEWITH SAID OTHER INPUT CIRCUIT, GAIN REGULATING MEANS CONNECTED WITH SAIDOUTPUT CIRCUIT AND IN CONTROLLING CONNECTION WITH SAID REGULATINGAMPLIFIER FOR CONTROLLING ITS GAIN IN DEPENDENCE UPON THE AMPLIFIEDOUTPUT VOLTAGE OF SAID DIFFERENTIAL AMPLIFIER AND IN THE SENSE TOWARDREDUCING SAID LATTER VOLTAGE, AN ANGLE TRANSMITTER CONNECTED WITH SAIDGAIN REGULATING MEANS TO BE CONTROLLED THEREBY AND HAVING AN OUTPUTSIGNAL INDICATIVE OF THE INSTANTANEOUS AMPLIFICATION GAIN OF SAIDREGULATING AMPLIFIER, AND A DIFFERENTIAL-FORMING NETWORK CONNECTEDBETWEEN SAID ANGLE TRANSMITTER AND SAID LONGITUDINAL CONTROL MOTOR FORCONTROLLING THE SAID LONGITUDINAL CONTROL MOTOR IN THE SENSE REQUIRED TOREDUCE SAID OUTPUT SIGNAL.